轴流压气机进口总压畸变的端区流动响应数值研究

To study the effect of the inlet circumferential total pressure distortion on the compressor performance, as well as the difference in the responses of end-wall flow to the inlet distortion under different distortion angles, annulus unsteady numerical simulations are performed on the first stage of a transonic axial compressor. A uniform inlet flow and three distorted flows with the distortion angles of 30°, 90° and 120° are adopted as the inlet conditions. Simulation results show that the inlet distortion significantly worsens the compressor performance, and the performance deterioration increases with the distortion angle. When the distortion angle is 120°, the structure and strength of the shock waves in the tip region of blades will change as the rotor blade sweeps through the distortion region, and the trajectory and strength of the tip clearance leakage vortexes also fluctuate periodically. The distortion also affects the flow in the downstream stator blades and aggravates the separation of the blade tip and corner. The number of the affected rotor passages reduces when the distortion angle decreases to 90°, but the dynamic response rule of the flow field to the distortion remains unchanged. However, the distortion completely decays after passing through rotor blades and the stator passages are almost unaffected when the distortion angel decreases to 30°. The results reveal the relationship of the change of compressor performance with the dynamic characteristics of flow structures at blade tip and root region and the distorted inlet flow, which provides a theoretical basis for improving the compressor's distortion resistance.